EP0360661A1 - Rotorcraft blade and rotor using same - Google Patents

Abstract

- According to the invention, said blade (1) is overtwisted in the end zone (8) from at least 0.85 R, in such a manner that the resultant twist is at least equal to a limiting twist theta limC dependent upon the shape in plan Co/C(r) of said blade in this zone and so that the coefficient of lift Cz has, in a first fraction of the end zone, extending at least between r = 0.85 R and r = 0.9 R, a decreasing value, remaining below a first limiting value Czlim1 = Czm - a(r/R - b), Czm being the mean coefficient of lift of said rotating wing system and a and b being experimental constants, and in a second end fraction extending between r = 0.9 R and at least r = 0.95 R, the coefficient of lift continues its decrease, remaining below or equal to a second limiting value Czlim2 which decreases linearly from the value Czlim1 for r = 0.9 R to the value zero for r = R. <IMAGE>

Description

The present invention relates to improvements to the blades of rotary wings for aircraft, as well as rotary wings comprising such improved blades.

We know that, if we consider a rotary blade, such that its successive elementary sections along the span have a constant cord and that the cords of said elementary sections are coplanar, the elementary forces of lift and drag linked to each of said elementary blade sections vary substantially as the square of the distance from an elementary section considered to the axis of rotation of the rotary wing. It follows that the end region of the blade has a significant influence on the aerodynamic functioning of the rotor and that the aerodynamic result of the lift and drag forces is approximately 75% of the span of the blade.

We also know that, to best adapt the incidence of successive elementary sections to the aerodynamic speeds they encounter due to the rotation of the blade, we generally proceed to twist such a blade around its longitudinal axis, in order to make it work with large setting in the vicinity of the axis of rotation of the rotary wing where the speed is low and with low setting towards the end of the blade where the speed is higher. Such a twist is generally linear, that is to say that the elementary variation dϑ of the twist angle ϑ over an elementary variation dr of span is constant. With such a measure, it is possible to improve the Cz / Cx fineness of operation of the profiles along the span of the blade.

However, as numerous experiments have shown, a very intense marginal vortex forms at the tip of the blade, resulting from the compensation of the pressure differences prevailing between the lower and upper surfaces of the sections close to the end. The presence of this marginal vortex leads on the one hand to an increase in the power consumed by the rotor and on the other hand to a significant emission of noise both at low forward speeds by interaction between the vortex and the following blades which l 'intercept, that at high forward speeds, due to the extreme aerodynamic operating conditions with the appearance of air compressibility under these conditions and the presence of shock waves.

The object of the present invention is a blade tip which significantly reduces the formation of the marginal vortex and which makes it possible to improve the aerodynamic performance of a rotary wing and to reduce its operating noise.

It consists in twisting the end zone of the blade from 70% of the span so that the coefficient of lift of the sections decreases progressively until a value substantially zero at the end of the blade. Thus, without affecting the overall lift of the blade end zone, the large pressure differences between the lower and upper surfaces are eliminated and consequently the conditions for the formation of the marginal vortex are eliminated.

(Czlim-Cz),
équation dans laquelle k = dCz/di représente la variation du coefficient de portance Cz par rapport à la variation d'incidence i en degrés, (pour un nombre de mach de 0,6 vers l'extrémité de pales, le coefficient k est alors égal à 0,13) et Czlim est une valeur limite supérieure du coefficient de portance Cz telle que, d'une part, dans une première fraction de la zone d'extrémité s'étendant au moins entre r = 0,85 R et approximativement r = 0,9 R, ce coefficient de portance limite a une valeur
(2) Czlim1 = Czm-a(r/R - b),
équation dans laquelle Czm est le coefficient de portance moyen de ladite voilure tournante, et a,b sont des constantes, au moins égales à 1,6 et 0,87 respectivement, et telle que, d'autre part, dans une seconde fraction de la zone d'extrémité s'étendant entre approximativement r = 0,9R et au moins r = 0,95 R, le coefficient de portance limite Czlim2 a une valeur définie par un segment de droite passant par la valeur de Czlim1 donnée par l'équation (2) pour r = 0,9 R et par la valeur Cz = 0 pour r = R.To this end, according to the invention, the blade for rotary wing of an aircraft, the successive elementary sections of which along the span R have a constant cord Co and are twisted from the root of the blade towards the end thereof at an angle ϑ whose variation d ϑ / dr as a function of the relative span r of an elementary section considered is constant, is remarkable in that said blade presents in the area end extending at least between 0.85 R and at least 0.95 R an overcasting such as:
- that the resulting twist is in said zone less than or at most equal, in absolute value, to a limit twist given by
(1) ϑlim = ϑ + $\frac{\text{1}}{\text{k}}$

(Czlim-Cz),
equation in which k = dCz / di represents the variation of the coefficient of lift Cz compared to the variation of incidence i in degrees, (for a number of mach of 0.6 towards the end of blades, the coefficient k is then equal to 0.13) and Czlim is an upper limit value of the lift coefficient Cz such that, on the one hand, in a first fraction of the end zone extending at least between r = 0.85 R and approximately r = 0.9 R, this limit lift coefficient has a value
(2) Czlim1 = Czm-a (r / R - b),
equation in which Czm is the average lift coefficient of said rotary wing, and a, b are constants, at least equal to 1.6 and 0.87 respectively, and such that, on the other hand, in a second fraction of the end zone extending between approximately r = 0.9R and at least r = 0.95 R, the limiting coefficient of lift Czlim2 has a value defined by a straight line passing through the value of Czlim1 given by equation (2) for r = 0.9 R and by the value Cz = 0 for r = R.

The first end zone fraction can extend in span between r = 0.7 R and approximately r = 0.9 R. In addition, said second end zone fraction can extend between approximately r = 0 , 9 R and r = R.

As will be shown hereinafter, if a rotary wing blade satisfies, with regard to its twist, under the above conditions, its distribution of lift coefficient in span, characterized by a progressive decrease to a zero value, makes it possible to substantially reduce the formation of the marginal vortex and therefore leads to an increase in performance and a decrease in operating noise.

It will be noted that the implementation of the present invention for the improvement of blades with linear twist is particularly advantageous, since the dynamic centering in rope of said blade in the end zone is not modified, so that the equilibrium general of the blade in operation on the aircraft is not altered.

In the case where, in said first and second fractions of the end zone, the chord of said successive elementary sections of the blade is no longer constant or quasi-constant, but evolves as a function of the position r in span of said elementary sections, the bearing surface of said blade at the end undergoes, with respect to a rectangular blade, a variation in bearing surface and therefore in lift coefficient. It is therefore necessary to take this into account.

(Czlim.Co/C(r) - Cz), expression dans laquelle le coefficient de portance Czlim est choisi, dans lesdites fractions de zone d'extrémité, au plus égal à une valeur limite CzlimC telle que
(4) CzlimC = Czlim.Co/C(r), expression dans laquelle Czlim représente soit Czlim1, soit Czlim2 définies ci-dessus pour une pale rectangulaire en fonction de celle desdites première ou seconde fractions de zone d'extrémité qui est concernée.Also, when in said first and second fractions of the end zone, said sections successive elementaries have an evolving chord C , which is a function C (r) of the position r in span of an elementary section considered, the twist of said blade is, in absolute value, advantageously at least equal to the limit twist ϑlimC given by
(3) ϑlimC = ϑ + $\frac{\text{1}}{\text{k}}$

(Czlim.Co/C(r) - Cz), expression in which the lift coefficient Czlim is chosen, in said end zone fractions, at most equal to a limit value CzlimC such that
(4) CzlimC = Czlim.Co/C(r), expression in which Czlim represents either Czlim1 or Czlim2 defined above for a rectangular blade as a function of that of said first or second end zone fractions which is concerned.

Experience has shown that the twist limits ϑlim and ϑlimC respectively given by equations (1) and (3) could possibly be exceeded, provided that the absolute value of the actual twist angle communicated to the blade is, in any point of said first and second zone fractions, at least equal to the corresponding value of ϑlim or ϑlimC, reduced by 0.5 °.

Furthermore, for construction reasons, it is often advantageous for said equations (1) and (3) to be applied in a second fraction of the end zone defined by the domain extending from approximately r = 0.9R to r = 0.95R and via the point r = R, a linear twist connection is made between r = 0.95R and r = R, whatever the function C (r).

• La figure 1 est une vue en plan schématique d'une pale rectangulaire pour voilure tournante.
• La figure 2 est un diagramme sur lequel sont représen­tées, en fonction de l'envergure relative r/R, la répartition réelle du coefficient de portance de la pale de la figure 1, et la répartition limite conforme à la présente invention.
• La figure 3 illustre la répartition limite d'angle de vrillage permettant à la pale de la figure 1 de satis­faire aux conditions de perfectionnement définies par la présente invention.
• La figure 4 est une vue en plan schématique d'une pale pour voilure tournante, à extrémité effilée de forme trapézoïdale.
• La figure 5 est un diagramme donnant, en fonction de l'envergure relative r/R, la limite de vrillage conforme à l'invention, pour la pale de la figure 4.
• La figure 6 montre, à plus grande échelle, l'extrémité d'une pale à bord d'attaque parabolique en extrémité.
• La figure 7 est un diagramme sur lequel est représentée l'évolution du vrillage de la pale montrée par la figure 6.
• La figure 8 est un diagramme montrant le vrillage conforme à l'invention de la pale de la figure 6, par rapport aux limites conformes à la présente invention.
• La figure 9 illustre schématiquement une installation pour l'essai de voilures tournantes permettant la mesure de leur portance et de la puissance consommée pour leur entraînement en rotation, ainsi que la mesure du bruit engendré.
• La figure 10 est un diagramme illustrant l'amélioration des performances aérodynamiques d'une voilure tournante équipée de pales telles que celle montrée par la figure 6 et survrillée selon l'invention, par rapport à une voilure tournante connue équipée de pales identiques mais à vrillage linéaire.
• La figure 11 est un diagramme illustrant la diminution du bruit de fonctionnement de la voilure tournante conforme à l'invention, par rapport à une voilure tournante identique, dont les pales sont vrillées linéairement.

The figures of the appended drawing will make it clear how the invention can be implemented. In these figures, identical references designate similar elements.

• Figure 1 is a schematic plan view of a rectangular blade for rotary wing.
• FIG. 2 is a diagram on which are represented, as a function of the relative span r / R, the actual distribution of the lift coefficient of the blade of FIG. 1, and the limit distribution according to the present invention.
• Figure 3 illustrates the limit twist angle distribution allowing the blade of Figure 1 to meet the refinement conditions defined by the present invention.
• Figure 4 is a schematic plan view of a blade for rotary wing, with a tapered end of trapezoidal shape.
• FIG. 5 is a diagram giving, as a function of the relative span r / R, the twist limit according to the invention, for the blade of FIG. 4.
• FIG. 6 shows, on a larger scale, the end of a blade with a parabolic leading edge at the end.
• FIG. 7 is a diagram on which the evolution of the twist of the blade shown in FIG. 6 is represented.
• Figure 8 is a diagram showing the twist according to the invention of the blade of Figure 6, with respect to the limits according to the present invention.
• FIG. 9 schematically illustrates an installation for testing rotary wings allowing the measurement of their lift and the power consumed for their rotary drive, as well as the measurement of the noise generated.
• FIG. 10 is a diagram illustrating the improvement in the aerodynamic performance of a rotary wing equipped with blades such as that shown in FIG. 6 and over-woven according to the invention, compared with a known rotary wing equipped with identical but twisted blades linear.
• FIG. 11 is a diagram illustrating the reduction in the operating noise of the rotary wing according to the invention, compared with an identical rotary wing, the blades of which are twisted linearly.

The blade 1, shown schematically in plan in Figure 1, is intended to be rotated counterclockwise (arrow F) around an axis 2, orthogonal to the plane of the figure, by l 'through a hub not shown. It has a blade root 3 allowing it to be connected to said hub. It is part of the rotary wing of an aircraft comprising a plurality of such blades 1.

The leading edge 4 and the trailing edge 5 of said blade 1 are parallel, so that, in plan, said blade is rectangular and the string of successive profiles constituting it is constant and equal to Co. The span of said blade, counted from said axis of rotation 2, is equal to R. We denote by r the span distance separating a particular section 6 from axis 2.

Furthermore, said blade is twisted around its twist axis 7, which can also be its pitch control axis. By "twisting" is meant the angular evolution along the span R of the zero lift rope of the profiles or sections 6.

Usually for helicopter blades, the twist of the blade 1 is constant along the span R, that is to say that it is a linear function of r . It is for example equal to - 10 ° (the minus sign designating that the setting of the profiles is decreasing from the root of the blade 1 towards its end).

If, for example by measuring the pressures exerted at different points distributed on said blade, the coefficient of lift Cz is determined along the latter, in its outer terminal part 8 comprised between the sections respectively distant from the axis 2 of 0.7 R distance and R distance, we can draw a curve whose shape is that of curve 9 shown schematically in Figure 2, for the flight domain of low and medium speeds, in level or low slope.

dans laquelle Fz est la portance de la voilure tournan­te, ρ la masse volumique de l'air, S la surface du disque formé par la voilure tournante ; σ la plénitude de ce disque (c'est-à-dire le rapport entre le produit du nombre de pales 1 par la surface d'une pale et la surface S) et U la vitesse relative de l'air ; pour la plupart ces hélicoptères, Czm est de l'ordre de 0,4 ou 0,5 ;
. λ est le paramètre d'avancement de l'aéronef, c'est-­à-dire le rapport entre la vitesse d'avancement dudit aéronef et la vitesse périphérique de son rotor ; pour la plupart des hélicoptères, λ est compris entre 0,25 et 0,35, pour les vitesses de croisière courantes ;
. Ψ est l'azimut de la pale 1 ( Ψ = 0 lorsque la pale 1 occupe la position arrière). Pour simplifier, on peut prendre un azimut moyen de 180°. Dans ces conditions, sin Ψ = 0 et λsin Ψ = 0.
. k1 et k2 sont des constantes voisines de 1 ; par exemple k1 est égal à 1,2 et k2 à 0,83 dans le cas particulier expérimental sur lequel la Demanderesse a réalisé des mesures de pression en vol.
- une seconde partie BC, également sensiblement linéai­re, correspondant à une seconde fraction de zone 8.2 de la partie terminale 8 et s'étendant entre des envergures relatives r/R respectivement égales à environ 0,9 et 1, et dans laquelle la valeur de Cz reste pratiquement constante ; et
- une troisième partie CD correspondant à un effondre­ment brutal du coefficient Cz au voisinage de l'enver­gure relative égale à 1.This schematic curve 9 essentially comprises three successive parts:
a first part AB, substantially linear and decreasing corresponding to a first fraction of zone 8.1 of the terminal part 8 and extending between relative spans r / R respectively equal to about 0.7 and 0.9, said first part AB possibly being represented by the function:
Cz = Czm - k1 (r / R-k2) - λ sin Ψ
in which :
. Czm is the mean lift coefficient of the rotary wing to which the blade 1 belongs, this mean lift coefficient Czm being itself defined by the expression

in which Fz is the lift of the rotary wing, ρ the density of the air, S the surface of the disc formed by the rotary wing; σ the fullness of this disc (that is to say the ratio between the product of the number of blades 1 by the surface of a blade and the surface S) and U the relative speed of the air; for most of these helicopters, Czm is of the order of 0.4 or 0.5;
. λ is the advancement parameter of the aircraft, that is to say the ratio between the advancement speed of said aircraft and the peripheral speed of its rotor; for most helicopters, λ is between 0.25 and 0.35, for current cruising speeds;
. Ψ is the azimuth of blade 1 (Ψ = 0 when blade 1 occupies the rear position). To simplify, we can take an average azimuth of 180 °. Under these conditions, sin Ψ = 0 and λsin Ψ = 0.
. k1 and k2 are constants close to 1; for example k1 is equal to 1.2 and k2 is 0.83 in the case experimental individual on which the Applicant carried out pressure measurements in flight.
a second part BC, also substantially linear, corresponding to a second fraction of zone 8.2 of the terminal part 8 and extending between relative spans r / R respectively equal to approximately 0.9 and 1, and in which the value of Cz remains practically constant; and
- a third part CD corresponding to a sudden collapse of the coefficient Cz in the vicinity of the relative span equal to 1.

Such a curve 9 is qualitatively representative of the evolution of the lift coefficient of all the rectangular blades with linear twist, in the usual case where said twist is kept constant until the end of the blade.

Because of the distribution of the lift coefficient Cz along the line BCD at the end of the blade, the marginal vortices are large, so that the performance of rotary wings equipped with such blades 1 are affected and the operation is noisy.

To remedy these drawbacks, the present invention provides, for example, for replacing, between the relative spans 0.9 and 1, the line BCD by a line FD, corresponding to a linear decrease to zero of the coefficient of lift Cz, the slope of the FD line being determined by the performance improvement sought. Thus, marginal eddies are minimized. In so doing, this results in a loss of lift for the blade corresponding to the area of the triangle 10 delimited between the lines BC, CD and FD.

Also, if one wishes to keep the overall lift constant, it is necessary to increase the lift coefficient in front of 0.9 R, for example between 0.7 R and 0.9 R.

This is obtained by a distribution of the coefficient of lift Cz, between 0.7 R and 0.9 R, such as that represented by the line EF, arranged above the line AB. Thus, between the lines AB and EF is delimited an area 11 corresponding to an increase in lift, capable of compensating for the loss represented by the triangle 10.

We therefore see that thus the distribution of the lift coefficient between r = 0.7R and r = R is, in accordance with the invention, determined by a curve 12, composed of the lines EF and FD. This curve 12 constitutes an upper limit, below which the actual distribution of the lift coefficient of the blade 1 must be found. The curve in phantom 13 gives an example of the actual distribution of lift coefficient.

In addition, according to the invention, the ordinate Y of the portion EF of the curve 12 satisfies the relation
Y = Czm - 1.6 (r / R - 0.87) - λsinΨ
in which Czm is the average lift coefficient of the rotary airfoil to which the blade 1 belongs, λ is the advance parameter of the aircraft and Ψ is the azimuth of the blade 1. For simplification, we can as previously take Ψ = 180 °, which leads to cancel λsinΨ. This portion EF of the curve 12 therefore constitutes an upper limit Czlim1 to the coefficient of lift Cz of the blade 1 improved according to the invention between 0.7 R and 0.9 R. Similarly, the portion FD of the curve 12, preferably straight, constitutes an upper limit Czlim2 to the lift coefficient Cz of the blade 1 improved according to the invention between 0.9R and R.

(Czlim - Cz)
dans laquelle k représente dCz/di (i = incidence). Généralement, k est voisin de 0,13.According to another feature of the present invention, to determine the curve 12 defining the limits Czlim1 and Czlim2, a lower limit is imposed on the absolute value of the twist of the blade 1, such as that represented by curve 14 in FIG. 3 In this FIG. 3, the relative span r / R of the different successive sections of the blade 1 comprised between r = 0.7R and r = R is shown on the abscissa and the difference Δ (in degrees) between the limit twist angle ϑlim of said sections and twist angle ϑ0.7 R of the section arranged at r = 0.7R. This curve 14 comprises a first part GH, corresponding to the part EF (Czlim1) of the curve 12, and a second part HI corresponding to the part FD (Cslim2) of the said curve 12. In FIG. 3, there is also shown a curve 15 defining a twist law corresponding to the distribution of the lift coefficient Cz, illustrated by curve 13 in FIG. 2, as well as two curves 16 and 17 corresponding to blades known for helicopters. Curve 16 relates to a rectangular blade having an overall linear twist of -7.3 ° between r = 0 and r = 0.925R and an unchanged twist between r = 0.925R and r = R. Curve 17 relates to a blade having a twist global linear of -12 ° between r = 0 and r = R. We see that these two curves 16 and 17 are arranged, totally or partially, outside the twist limit ϑlim defined by curve 14. This twist limit ϑlim is given by the expression
(2) ϑlim = ϑ + $\frac{\text{1}}{\text{k}}$

(Czlim - Cz)
in which k represents dCz / di (i = incidence). Generally, k is close to 0.13.

(Czlim - Cz), dans lequel
Czlim est la valeur définie par la courbe 12 pour ladite distance r et Cz est la valeur du coefficient de por­tance pour ladite section 6 avant perfectionnement conformément à l'invention.Thus, for any elementary section 6 of the blade 1, whose span distance r is between 0.7R and R, the limit twist angle ϑlim according to the improvement of the invention is equal to the twist angle ϑ of said section before improvement, increased by the corrector term $\frac{\text{1}}{\text{k}}$

(Czlim - Cz), in which
Czlim is the value defined by curve 12 for said distance r and Cz is the value of the lift coefficient for said section 6 before improvement in accordance with the invention.

- la ligne (III) indique une répartition désirée pour le coefficient de portance limite Czlim (Czlim1 et Czlim2) conforme à la courbe 12 de la figure 2 ;
- la ligne (VI) indique la répartition de l'angle de vrillage de l'équation (1) ci-dessus, les lignes (IV) et (V) présentant les calculs intermédiaires de Czlim-Cz et de

- la ligne VII donne la différence entre l'angle de vrillage limite à une position r et l'angle de vrillage ϑ 0,7R à r = 0,7R (courbe 14 de la figure 3).In Table I below, a numerical example is given allowing a good understanding of the present invention. This table I comprises seven lines I to VII, each containing nine values each corresponding to a relative span value r / R, said relative span values being staggered between 0.7 and 1. It relates to the improvement according to invention of a rectangular blade 1, the twist of which conforms to that given by curve 16 of FIG. 3:
- Line (I) indicates the twist of the rectangular blade 1 before overwiring according to the invention. We can see that said blade is twisted linearly at an angle of -7.3 ° up to r / R = 0.925, said twisting angle remaining constant between r / R = 0.925 and r / R = 1 (see curve 16 );
- Line (II) indicates the measured distribution of the lift coefficient Cz in wingspan before the over-rolling of the invention, this distribution being in accordance with curve 9 of FIG. 2;

- the line (III) indicates a desired distribution for the limiting coefficient of lift Czlim (Czlim1 and Czlim2) in accordance with curve 12 of FIG. 2;
- line (VI) indicates the distribution of the twist angle of equation (1) above, lines (IV) and (V) presenting the intermediate calculations of Czlim-Cz and

- Line VII gives the difference between the twist angle limited to a position r and the twist angle ϑ 0.7R to r = 0.7R (curve 14 in Figure 3).

In Figure 4, there is shown a blade 1.1 in all respects identical to the blade 1 shown in Figure 1, except as regards the end part. In fact, instead of being rectangular to its end like the blade 1, the blade 1.1 is tapered at the end, because the trailing edge 5 has an inclined end portion 5.1 approaching the leading edge 4. This inclined end portion 5.1 of the trailing edge 5 is rectilinear and it begins at r = 0.75R, where the profile chord is equal to Co, to end at r = R, where the profile chord is no longer then than Co / 3.

Thus, in the blade 1.1, the bearing surface decreases from r = 0.75R to r = R, so that this results in a reduction in lift favoring the action of the overcasting according to the present invention. It is therefore necessary to take account of this variation in bearing surface, when it is desired to apply to a blade with an evolving profile, the limits defined above for a rectangular blade.

Generally, if C (r) represents the evolution of the chord C as a function of the distance r , the Czlim limits (Czlim1 and Czlim2) defined above for a rectangular blade are corrected, so as to define a new limit twist angle vlimC and a new lift coefficient limit Czlim1, as defined by equations (3) and (4), above.

Thus, it is possible to determine, at each point of the end 8 of the blade 1.1, the limit twist angle according to the present invention, from the characteristics of the rectangular blade of FIG. 1.

In Table II, a numerical example is given, to be compared to the numerical example in Table I.

In Table II,
- lines (I), (II) and (III), are identical to the corresponding lines in Table I and indicate respectively the twist, the measured distribution of the lift coefficient and the desired distribution of the lift coefficient of the rectangular blade 1 before improvement according to the invention;
- the line (IV) indicates the Co / C ratio (r), representative of the tapering of the end of the blade 1.1 of FIG. 4;
- line (V) gives the limit ϑlimC calculated by equation (4) above; and
- line (VI) gives the difference Δ C = ϑ limC - ϑ 0.7R between the calculated limit ϑlimC and the twist angle ϑ0.7R at distance r = 0.7R. TABLE II r / R 0.70 0.75 0.80 0.85 0.90 0.925 0.95 0.975 1 ϑ (I) -5.12 -5.48 -5.85 -6.21 -6.58 -6.76 -6.76 -6.76 -6.76 Cz (II) 0.56 0.50 0.44 0.38 0.32 0.32 0.32 0.32 0.32 Czlim (III) 0.68 0.60 0.52 0.44 0.36 0.27 0.18 0.09 0 Co / C (r) (IV) 1 1 1,154 1.364 1.667 1.875 2.143 2.5 3 ϑlimC (V) -4.20 -4.71 -4.62 -4.52 -4.43 -5.33 -6.25 -7.49 -9.22 ϑlimC - ϑ0.7R (VI) 0 -0.51 -0.42 -0.32 -0.23 -1.13 -2.05 -3.29 -5.02

In FIG. 5, the difference ΔC (in degrees) has been represented as a function of the relative span r / R and the corresponding curve 18 has been obtained.

In FIG. 6, related to the abscissa axis r / R of FIG. 7, the end of a blade 1.2 is shown. This blade 1.2 is identical in all respects to blade 1 in FIG. 1, except as regards the end of the leading edge 4. In fact, between r = 0.9475R and r = R, instead of be straight and parallel to the trailing edge 5, the leading edge 4 comprises a parabolic part 4.1, for example such that
C (r) / Co = 1 - 234.375 (r / R - 0.9466) ²
At r = 0.9475R, the chord of the blade is equal to Co, but at r = R this chord is equal to Co / 3.

From the above, it appears that one can easily calculate the limit ϑlimC corresponding to said blade 1.2. The corresponding calculations are collated in Table III, below. This limit ϑlimC, or rather the difference ΔC between it and the twist angle ϑ 0.7R at r = 0.7R is represented by curve 19 in Figure 8. To draw this curve 19, between r = 0 , 7R and r = 0.95R, as well as for r = R, the results given in Table III were used. On the other hand, between r = 0.95R and r = R, a 19.1 connection with linear twist was made.

In FIG. 8, from curve 19 which is representative of a twist limit which can be described as "theoretical", a curve 20 has been drawn representing Δ C + 0.5 °, this value 0.5 ° constituting a tolerance with respect to the limit defined by curve 19. We thus obtained the "practical" limit of TABLE III r / R 0.70 0.80 0.85 0.90 0.95 1 ϑ (I) -5.12 -5.85 -6.21 -6.56 -6.76 -6.76 Cz (II) 0.56 0.44 0.38 0.32 0.32 0.32 Czlim (III) 0.68 0.52 0.44 0.36 0.18 0 Co / C (r) (IV) 1 1 1 1 1 3 ϑlimC (V) -4.20 -5.23 -5.75 -6.25 -7.84 -9.22 ϑlimC - ϑ0.7R (VI) 0 -1.03 -1.55 -2.05 -3.64 -5.02 twist illustrated by curve 20. This practical twist limit made it possible to define the real twist, in accordance with the invention, of the blade 1.2. This real twist is represented by curve 21 in FIG. 7 and illustrated by means of the difference Δ C, by curve 22 in FIG. 8.

In FIGS. 7 and 8, curves 23 and 24 are also shown, illustrating a linear twist of -10 ° for the blade 1.2. It can be seen that the vane 1.2 overcast according to the invention is twisted linearly between r = 0.7R and r = 0.85R, then overcast as described above between r = 0.85R and r = R.

To illustrate the improvements brought by the present invention, comparative tests were carried out between a rotary wing 25.1 equipped with three blades 1.2 linearly twisted (in accordance with curves 23 and 24) and an identical rotary wing 25.2 equipped with three blades 1.2 swiveled at the end according to the present invention (in accordance with curves 21 and 22).

These comparative tests were carried out using the bench 26 shown schematically in FIG. 9.

This test bench 26 essentially comprises a balance 27, articulated at 28 and carrying a motor 29, a transmission 30 and a hub 31, to which can be attached either the rotary wing 25.1 or the rotary wing 25.2.

Said hub 31 is arranged at one end of the balance 27, while a dynamometer 32 is connected to the other end of said balance. It is understood that when a rotary wing is fixed to the hub 31 and is rotated by the motor 29 and the transmission 30, of which the power delivered can be measured with precision by means not shown, it exerts a thrust Fz capable of being measured by the dynamometer 32.

In addition, the test bench 26 includes microphones 34 connected to an apparatus 33 for measuring sound level.

Various tests were carried out by successively fixing the rotary wings 24.1 and 24.2 to the hub 31 and the powers absorbed by these wings were measured as a function of the reduced thrust. z for three peripheral speeds Up, respectively equal to 210 m / s, 225 m / s and 235 m / s. The results are illustrated in FIG. 10, in which DW represents, as a percentage, the reduction in power absorbed by the rotor 25.2 compared to the power absorbed by the rotor 25.1 to ensure a given thrust. Such a reduction in absorbed power is of the order of 3 to 4% for the levitation and normal peripheral speeds. It can reach 10% for strong pushes ( z = 24) and the peripheral speed of 235 m / s. The aerodynamic performance of the wing according to the invention is therefore greatly improved compared to that of known wings.

Furthermore, in FIG. 11, the noise reduction ΔN (in dB) of the rotary wing 25.2 according to the invention is shown, relative to the wing 25.1. FIG. 11 shows that, in hovering flight and for the usual thrusts ( z = 15), a noise reduction of the order of 2dB is obtained by the rotary wing 25.2 compared to the rotary wing 25.1, for a peripheral speed range between 220 and 230 m / s. In forward flight, for high-speed phases or for descent phases, where marginal blade-vortex interactions have more of opportunities to occur, greater reductions in noise are seen.

Claims (8)

1. Blade for rotary wing of an aircraft, the successive elementary sections of which along at least most of the span R have a constant chord Co and are twisted from the root of said blade towards the end of this of an angle ϑ whose variation dϑ / dr as a function of the position r in span of an elementary section considered is constant, characterized in that said blade present in the end zone extending at least between 0.85 R and at least 0.95 R an overshoot such that the resulting twist is in said zone, less than or at most equal, in absolute value, to a limit twist given by
(1) ϑlim = ϑ + $\frac{\text{1}}{\text{k}}$

(Czlim - Cz) equation in which k is the ratio of the variation of the lift coefficient Cz to the variation of incidence i and Czlim is an upper limit value of the lift coefficient Cz such that, on the one hand, in a first fraction of the end zone extending at least between r = 0.85 R and approximately r = 0.9 R the limit lift coefficient has a value
(2) Czlim1 = Czm - a (r / R - b), equation in which Czm is the average lift coefficient of said rotary wing, a and b constants at least equal to 1.6 and 0.87 respectively and such that, on the other hand, in a second fraction of the end zone extending between approximately r = 0.9 R and at least r = 0.95 R, the limiting coefficient of lift Czlim2 has a value decreasing linearly from the value of Czlim1 given by equation (2) for r = 0.9R up to zero for r = R. 2 - Blade for rotary wing according to claim 1,
characterized in that said first end zone fraction extends between r = 0.7R and approximately r = 0.9R. 3 - Blade for rotary wing according to one of claims 1 or 2,
characterized in that said second end zone fraction extends between approximately r = 0.9R and r = R. 4 - Blade for rotary wing according to any one of claims 1 to 3, such that in said first and second end zone fractions, said successive elementary sections have an evolving chord C which is a function C (r) of the position r in span of an elementary section considered,
characterized in that, over the extent of said first and second end zone fractions, the twist of said blade is, in absolute value, advantageously at least equal to the limit twist ϑlimC given by
(3) ϑlimC = ϑ + $\frac{\text{1}}{\text{k}}$

(Czlim.Co/C(r) - Cz)
equation in which the Czlim lift coefficient is chosen from said end zone fractions at most equal to a CzlimC value such that
(4) CzlimC = Czlim.Co/C(r), expression in which Czlim represents either Czlim1 or Czlim2 depending on which of said first or second fraction of end zone is considered. 5 - Blade for rotary wing according to any one of claims 1 or 4,
characterized in that, over the extent of said first and second end zone fractions, the twist of said blade is, in absolute value, at least equal to said limit twist less a tolerance. 6 - Blade according to claim 5,
characterized in that said tolerance is of the order of 0.5 °. 7 - Blade for rotary wing according to any one of claims 1, 4, 5 or 6,
characterized in that said second fraction of the end zone is defined by the range extending from approximately r = 0.9R to r = 0.95R and by the point r = R, a connection with a linear twist being produced between r = 0.95R and r = R, whatever the function C (r). 8 - Rotary wing for aircraft,
characterized in that it is provided with blades as specified according to any one of claims 1 to 7.

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